(1) Field of the Invention
The invention relates to a T-profile junction of composite materials. The junctions are used in the field of airplane airframe composite structures.
(2) Background Art
Aircrafts of 21st century make use of composite materials in their structural design as much as possible. But there are limitations due to the nature of composite materials making it impossible to use composite parts at some occasions. The so called “T-Pull problem”, where T refers to the profile of the composite part, is one of the above mentioned limitations, and because of the T-Pull problem some composite parts have to be designed unnecessarily heavy or in some cases have to be replaced by metallic parts making the parts even heavier.
The T-Pull problem is caused by the lack of “out of plane capabilities” of common composite materials. Composites differ from metals as their properties are not uniform in all directions. Especially the short transverse strength of composite materials is poor. For parts loaded in a direction out of plane the thickness stresses are inevitable and metallic materials are superior to composites for said applications. For the above reason, when using composite materials, conventional metallic designs of structures are altered to avoid out of plane stresses in the structure. But there are still cases for which changing the design geometry to avoid out of plane stresses conflict with functional requirements of the structure. Whenever a T/Junction exists in a composite part (I section, T section), the pulling load and the resulting resisting forces at flanges of a T-profile junction cause out of plane stresses/strains in the composite structure around the T-profile junction. These stresses can not be handled by composite materials, hence leading to the failure of the junction. The failure generally shows itself as delamination between C shaped web ply packages and any triangular filler at the T-profile junction. Delamination can also occur between lower cap ply package and filler as well. In both of the above cases, delamination continues between the web ply packages and cap ply packages, splitting the T section into 2 C section parts connected at the web. Those C sections, which encounter extreme deformation, against which they are not sized for, also develop delamination within the laminate leading to complete failure of the part.
The problem can not be solved by simply adding extra composite material. The amount of extra plies to be added usually would be as much that the composite material can be replaced by a metal. Said solution is absolutely ineffective, since the amount of composite material to be added is not linearly linked to the increase in stiffness. Even for some cases adding as much as possible composite material in given space does not solve the problem.
Consequently some previously developed composite aircraft structures with T-profile junctions are supported by metallic parts, when possible, such supports being L shaped brackets or square tension bars to support the flanges as shown in FIG. 1 of the attached drawings. If the space given is limited, brackets or tension bars can not be used.
The document EP1194285B1 discloses a fiber composite material with reduced delamination, said composite itself being modified to make it more delamination resistant with the consequence that no conventional composite materials may be used.
The document U.S. Pat. No. 4,966,802 discloses fibre reinforced resin composites formed by elements joined by a high shear strength, high fracture toughness adhesive. The elements are created in easy to produce cross-sectional shapes such as flat, C-shaped, Z-shaped or T-shaped. When joint, the elements form delamination resistant fibre reinforced resin composites having more complex shapes. The strength of the junctions between a pair of elements, such as a panel and its associated reinforcing members, can be enhanced by fasteners, such as rivets, if desired. Said state of the art is not specific to structures with T or I cross-section which are subject to pull loads such as typical spars, ribs on composite wings, stringers on pressure bulkheads, frames on landing gear bays and frames on cargo & passenger doors.
The document EP2021164A1 is dealing with the problem of delamination when a laminated element has a curve formed therein. An apparatus is disclosed for and a method of compressing a curved region of a laminated element or structure so as to protect the element or structure against delamination. Said state of the art is applicable to single laminates with L or C sections, where it is possible to compress the laminate from both surfaces, namely at the corner where it is most subject to delamination and to counteract out of plane loads. Said state of the art is not applicable to T-sectioned composite parts. Said state of the art does not counteract the main loading on the part, which causes the delamination, but only counteracts internal loads.
The document U.S. Pat. No. 4,909,659 A discloses projections formed in a wing skin to mate with recesses formed in a support substructure which, when positioned together, form interdigitations which prevent sliding displacement of the wing skin. A plurality of sliding bars transversely pass through the interdigitating sections to maintain an interlocking relation. The transversely oriented bars enable the use with contoured wing skin structures.
The document US2006243860 A1 discloses a composite stringer and skin structures and methods for forming the same. In one embodiment, a composite stringer and skin structure includes a polymer-based elongated stringer portion having reinforcing fibers positioned in a plurality of adjacent plies, a first portion of the reinforcing fibers being oriented at a relatively shallow angle relative to a selected reference direction, and a second portion of the reinforcing fibers being oriented at a relatively broad angle relative to the selected reference direction. A polymer-based and fiber reinforced skin member adjoins the stringer portion, and an adhesive material is interposed between the stringer portion and the skin member.